2006_6153-T-50-A-50 High Angle of Attack Characteristics

American Institute of Aeronautics and Astronautics

1

T-50/A-50 High Angel of Attack Characteristics

Myung Sup Lee*, Young Ho Lee†, and Ilwoo Lee‡ Korea Aerospace Industries, LTD., Sacheon, Republic of Korea

This paper describes the T-50/A-50 high AoA characteristics and the successful completion of the high AoA flight test as a part of the full-scale development (FSD). This paper presents three major subjects: The design goals, the predictions by the wind tunnel test data and the T-50/A-50 high AoA characteristics obtained by the flight test. It is presented that the T-50/A-50 has excellent high AoA characteristics, recovery capabilities and extreme resistance to the departure by passing over the major tasks of the FSD progress.

Nomenclature FSD = full scale development AoA = angle of attack CG = center of gravity OFP = operational flight program MAC = mean aerodynamic chord LEF = leading edge flap TEF = trailing edge flap HT = horizontal tail SB = speed brake LCDP = lateral control departure parameter HQS = handling quality simulator MPO = manual pitch override SUU = suspended under-wing unit SRC = spin recovery chute FLCS = flight control system PA = power approach UA = up and away Cm = pitching moment coefficient CN = normal force coefficient Cn = yawing moment coefficient Cnβ = directional stability derivative for sideslip

I. Introduction he T-50/A-50 is an advanced jet trainer/light attacker, which was recently developed and just started its mass produce in Republic of Korea. The T-50/A-50 needed to satisfy the high AoA (Angle of Attack) characteristics

requirements from the initial design phase. The design goals and methodologies to satisfy those requirements are presented referring to the lessons learned from other programs, like the deep stall of F-16 or the falling leaf of F-18 etc.

Several wind tunnel tests were performed to gather the aerodynamic data of high AoA region, and the aerodynamic coefficients as well as the predictions from tested data are presented. The deep stall CG (Center of Gravity) condition was predicted and the configuration change effect was tested to prevent and insure recovery from deep stall. And the departure resistance is estimated based on the criteria from other reference.

The flight test was performed as a final verification of high AoA characteristics and the test results and comparisons to wind tunnel test are shown. The pitching moment was extracted by performing intentional departure

* Research Engineer, Flight Science Section, Aircraft R&D Division † Senior Research Engineer, Flight Science Section, Aircraft R&D Division ‡ Principal Research Engineer & Chief, Advanced Development Programs Section, Member AIAA.

T

AIAA Atmospheric Flight Mechanics Conference and Exhibit21-24 August 2006, Keystone, Colorado

AIAA 2006-6153

Copyright © 2006 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


2

maneuvers. And the major aircraft motion parameters from departure resistance maneuvers performed are also summarized.

II. Design Goals of T-50/A-50 High AoA Aerodynamic Characteristics Design goals of T-50/A-50 high AoA aerodynamic characteristics are the departure resistance, deep stall

prevention, and spin-proof.

A. Departure Resistance Aircraft should not depart by normal operation within the flight manual limits. The ultimate goal is to making

“Care-Free” aircraft without any performance degradation. To achieve this design goal, lateral/directional stability and control surface effectiveness were analyzed and appropriate limit logics were designed and applied to OFP (Operational Flight Program). Applied limiters to T-50/A-50 OFP are AoA-g limiter, rudder fader, roll rate limiter, high AoA departure prevention logic, and anti-spin logic. Besides these limiters, T-50/A-50 has aural warning system in departure possible situation.

B. Deep Stall Prevention The aircraft possibly departs by intention or improper energy management, even with above-mentioned limiters.

In this departure situation, the T-50/A-50 was required to self-recover or to recover with simple procedure. To comply this requirement, aircraft should have no deep stall trim condition or recovery capability from deep stall. The T-50/A-50 is predicted to have inverted deep stall at initial design phase, but deep stall characteristics improved by additional control surface effect survey.

C. Spin-Proof The aircraft was required not to spin. To satisfy this goal, rotary-balance and free spin wind tunnel test were

performed, and anti-spin logic was applied to OFP based on predicted spin mode. (The spin characteristics are important, but are not dealt in this article.)

III. T-50/A-50 High AoA Aerodynamic Characteristics Prediction (Wind Tunnel Test) In this section, the T-50/A-50 high AoA aerodynamic characteristics are presented base on wind tunnel test

result. Main purpose of wind tunnel test in stability and control respect is gathering aerodynamic coefficients and every control surface effects for simulation aerodynamics database. And high AoA aerodynamic coefficients are used to check deep stall existence in departure and to predict control surface effect on deep stall recovery.

All aerodynamic characteristics presented in this section are based on low speed and rotary balance wind tunnel test. Low speed wind tunnel test is shown in Figure 1 and rotary balance test is in Figure 2.

Figure 1. Low Speed Wind Tunnel Test


3

Figure 2. Rotary Balance Wind Tunnel Test

A. Deep Stall Deep stall called as ‘high AoA pitch hang-up’ is defined as

trimmed condition by stable slope of pitching moment in out-of-control with maximum recovery control surface deflection. (Figure 3)

Deep stall occurs when CG is behind deep stall CG. So, deep stall CG should be beyond operational CG range to prevent deep stall. Deep stall CG limit is determined at pitching moment equilibrium condition with design margin. There are several studies about the design margin, and Table 1 shows one of reasonable guideline1,2. For T-50/A-50 deep stall static prediction, -0.07rad/s2 design margin was used. Therefore, deep stall CG limit is conceptually determined as Figure 4.

B. Deep Stall CG Estimation The T-50/A-50 deep stall CG limit was calculated with full recovery control configuration, i.e. full nose down

for upright and full nose up for inverted. Figure 5 shows Cm vs. CN plots which include the comparison between T-50/A-50 deep stall CG limit and operational most aft CG. As the plots show, T-50/A-50 was predicted to have enough deep stall margin for upright, but to have inverted deep stall trim condition considering operational CG range.

BAR0104 TEST DATAERECT DEEP STALL C.G. LIMIT CALCULATION

CN

Cm

T-50Most Aft CG

UprightDeep StallCG Limit

Clean Configuration

BAR0104 TEST DATAINVERTED DEEP STALL C.G. LIMIT CALCULATION

CN

Cm

InvertedDeep StallCG Limit

T-50Most Aft CG

Clean Configuration

Deep Stall TrimPredicted

Figure 5. Upright/Inverted Deep Stall Limit

-5.0-0.07Safety

-24.0-0.25Desirable

Pitch RateIn 2 sec After Recovery

Initiation (deg/sec)

Pitch AccelerationIn 1 sec After Recovery

Initiation (rad/sec2)

-5.0-0.07Safety

-24.0-0.25Desirable

Pitch RateIn 2 sec After Recovery

Initiation (deg/sec)

Pitch AccelerationIn 1 sec After Recovery

Initiation (rad/sec2)

Table 1. Deep Stall Recovery Pitch Control Margin Guideline

Cm

+ δHT = -Max

δHT = 0

δHT = +Max

α

Deep StallTrim Point

-

xCm

+ δHT = -Max

δHT = 0

δHT = +Max

α

Deep StallTrim Point

-

x

Figure 3. Deep Stall

Figure 4. Deep Stall DG Limit EstimationCN

Cm

Max.Nose DownControl

∆Cm forPitch AccelerationRequirement

0


4

C. Low Speed vs. Rotary Balance Wind Tunnel Test Result Comparison Predicted deep stall CG’s using low speed test data and rotary balance test data are compared. As presented in

Figure 6, deep stall CG limit predicted by low speed data shows more favorable result. Low speed test result shows approximately +2.6/4.0%MAC CG margin for upright and inverted deep stall. Tunnel to tunnel difference of predicted deep stall is quite big, considering that configuration change option effect on deep stall is 1 to 2%MAC.

Between two different predictions, rotary balance test result was selected for further design and analysis in conservative manner considering additional CG shift of production aircraft. (For F-16 case, both low speed and rotary balance wind tunnel test were performed and showed difference result. Between two results, rotary balance test result was closer to flight test result than low speed test.)

INVERTED DEEP STALL C.G. LIMIT CALCULATION

CN

Cm

Rotary Balance W/T

Low Speed W/T

UprightInverted

Figure 6. Low Speed W/T Test vs. Rotary Balance W/T Test Results

D. Control Surface Change Options To Extend Inverted Deep Stall CG Several configuration options were tested in wind tunnel to improve/resolve the inverted deep stall after rotary

balance test data were decided to be used for deep stall CG limit prediction. An outer mold line (OML) change was excluded from resolution options by the program policy. Considered options were as follows;

1) Leading edge flap (LEF) effect: LEF down 2) Trailing edge flap (TEF) effect: TEF up 3) Horizontal tail (HT) effect: reduced deflection angle of nose-up 4) Speed brake (SB) open effect LEF and TEF deflection options were to induce pitching moment increase. Reduction of HT maximum

deflection was to check pitching moment change, i.e. pitching moment decrease in maximum HT deflection by local flow separation. And open SB option was expected to increase recovering pitching moment and effective HT control power. The test results are shown in Figure 7. The margins are the differences between calculated deep stall CG limits from tested options and most aft CG condition.

-3.4

-2.9

-2.2

-1.2

-1.6

+0.8

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

Aft CG Margin (%MAC) Against Inverted Deep Stall

TEF = -20 / SB = +56

Max AB Nozzle

SB = +56

TEF = -20

LEF = +30

LEF/TEF/HT // SB = -2/0/-30/0 // 0

Figure 7. Upright/Inverted Deep Stall Limit


5

As result shows, nozzle, SB, and TEF effects on inverted deep stall are apparent, showing -1.6%MAC, -1.2%MAC and -2.2%MAC margin with nozzle, SB, and TEF respectively (baseline: -3.4%MAC margin) even though LEF effect is small. But nozzle could not be used as recovery control cause of engine control logic change to fixed nozzle at high altitude. Reduction of HT maximum deflection was negligible.

Based on this test results, OFP was modified to automatically move TEF down at inverted departure situation and speed brake could be included in recovery procedure after flight test verification.

E. Deep Stall Recovery Simulation For the next step, off-line 6-DOF deep stall recovery was

simulated to validate static analysis and to check additional result related to deep stall recovery. The simulation conditions are summarized as follows:

1) T-50/A-50 clean configuration 2) CGIDS (inverted deep stall CG limit considering SB effect)

-1 to +3 %MAC 3) Inverted deep stall trim at 35kft (Instantaneous trim for

simulation) 4) Check self-recovery: Check pitch acceleration, pitch rate

and altitude loss 5) Check SB open recovery: Check pitch acceleration, pitch

rate and altitude loss Figure 8 shows the concept of simulation. Simulations were performed from CGIDS –1 to +3%MAC. CGIDS

is statically predicted inverted deep stall CG in section A. Simulation result is summarized in Figure 9. Deep stall did not exist at (CGIDS –1)%MAC and aircraft self-recovered. From CGIDS, aircraft did not self-recover and recovered with SB open. At (CGIDS +3%MAC), aircraft did not recover even with SB open. Therefore, in simulation, aircraft could recover with CG’s up to (CGIDS +2%MAC) with SB open. This discrepancy between static analysis and simulation is due to the pitch acceleration margin considered in static calculation. And that pitch acceleration functions as recovering power in simulation.

The pitch acceleration in 1 sec and pitch rate in 2sec after recovery initiation were compared to reference design margin shown in Table 1. As shown simulation results in Figure 10, the “safety” criteria (0.07rad/s2 pitch acceleration in 1sec and 0.05deg/s pitch rate in 2sec) are considered proper to estimate the recovery capability from departure comparing the result at CGIDS and (CGIDS +2%MAC) at which the aircraft did not self-recovered but recovered with SB open.

ALTITUDE LOSS(during Inverted Deep Stall Recovery)

Deep StallFree Recoverable

w/ SB

NotRecoverable

with SB

0

2000

4000

6000

8000

10000

CGIDS - 1

CGIDS - 1

CGIDS

CGIDS +

1

CGIDS +

2

CGIDS +

3 CG

Alti

tude

Los

s ∆

h, ft

MAXIMUM PITCH ACCELERATION(during Inverted Deep Stall Recovery)

Deep StallFree Recoverable

with SB

NotRecoverable

with SB

0.0

0.2

0.4

0.6

0.8

1.0

CGIDS - 1

CGIDS - 1

CGIDS

CGIDS +

1

CGIDS +

2

CGIDS +

3 CG

Max

Qb

dot,

rad/

s2

Figure 9. Deep Stall Recovery – Altitude Loss and Maximum Pitch Acceleration

Figure 8. Deep Stall Recovery Simulation


6

PITCH ACCELERATIONAt 1 second After Recovery Initiation

0.00

0.05

0.10

0.15

0.20

CGIDS - 1

CGIDS

CGIDS +

1

CGIDS +

2

CGIDS +

3

CG, %MAC

Qbd

ot, r

ad/s

2

SB ClosedSB Open

0.07 rad/s2

Filled: Not Recovered

PITCH RATEAt 2 seconds After Recovery Initiation

0.0

5.0

10.0

15.0

20.0

CGIDS - 1

CGIDS

CGIDS +

1

CGIDS +

2

CGIDS +

3

CG, %MAC

Qb,

deg

/s

SB ClosedSB Open

5.0 deg/s

Filled: Not Recovered

Figure 10. Deep Stall Recovery –Pitch Acceleration and Pitch Rate

F. Inertial Coupling Cross-coupling, i.e. pitch and yaw motion during

roll maneuver, was analyzed to estimate departure by cross-coupling induced pitch up motion. Cross coupling is affected by inertia and gets bigger in fuselage-loaded aircraft which has bigger Iyy than Ixx. Figure 11 shows the calculation of CG limit able to suppress additional pitching-up moment by inertial coupling. Additional pitching-up moment was calculated by following simplified equation.

cSqIII

PCref

xzzzxxstbm ××

×+××××=∆

αα 2cos2sin)(5.02

Low speed wind tunnel data instead of rotary balance data are used for inertial coupling analysis considering limit AoA usually coupling occurs. T-50/A-50 applies rudder fader and roll command limiter along with AoA-g limiter, so departure susceptibility is generally small. And prediction shows enough margin comparing to operational CG range.

G. Departure Resistance Evaluation There are several well-known departure resistance evaluation methods using aerodynamic coefficients and

derivatives; Cnβ dynamic and LCDP (Lateral Control Departure Parameter). For T-50/A-50’s departure resistance was estimated using these parameters and compared to the criteria in early design phase. Figure 12 show Cnβ dynamic and LCDP of T-50/A-50 clean configuration. And Weissman criteria combining LCDP and Cnβ dynamic are plotted in Figure 13. 1. Cnβ dynamic

ααβββ

sincos ××−×= lxx

zznn C

IICC

dyn

Design goal of T-50/A-50 was Cnβ dynamic > 0.04 including margin and low speed wind tunnel data satisfied that criteria over operational AoA range. 2. LCDP

0cos ⟩×−×=a

a

l

nln C

CCCLCDP

δ

δ

ββα

CN

Cm

Inertia Coupling

Max. nose Down Control

Figure 11. Inertial Coupling CG Limit Estimation


7

Negative LCDP is known possible to generate roll reversal. Both low speed and rotary balance wind tunnel test result are presented in Figure 12. The results are open-loop characteristics not to include the augmentation by flight control logic. As the plot shows, LCDP remains positive over operational AoA range.

Cn_Beta_DynamicLEF = Scheduled, TEF = 0, HT = 0

AOA, Deg

Cn β

dyna

mic, D

eg

LCDP

AOA, Deg

LCD

P

Low Speed W/TRotary Balance W/T

Figure 12. Cnβ dynamic and LCDP for T-50 Clean Configuration

3. Weissman Criteria Weissman suggested Weissman Criteria1,3,4,5, correlated Cnβ dynamic and LCDP, to estimate departure

resistance. T-50/A-50 low speed and rotary balance wind tunnel test result are shown with Weissman criteria in Figure 13. The analysis is open-loop characteristics same as LCDP, and the result would change with augmentation by control law such as aileron-rudder interconnection.

As AoA increase over operational range, data go to ‘Mild Departure’ range similar to trend of LCDP. And rotary balance test data show worse characteristics than low speed data.

Cn Beta Dynamic

LCD

P

Rotary Balance W/TLow Speed W/T

No Departures

Poor RollControl

MildDepartures

Dep

artu

res/

Spin

-Pro

ne

Figure 13. Weissman Criteria for T-50 Clean Configuration

Based on these results, T-50/A-50 was deemed as departure resistant within limit AoA considering the applied limiting logics in OFP.

IV. T-50/A-50 High AoA Flight Test It was verified that the design goals, i.e. departure resistance, deep stall prevention, and spin-proof, were

satisfied through wind tunnel test data analysis, predictions, and off-line simulations. Then, T-50/A-50 flight test was performed as a final verification of high AoA characteristics. FSD fight test started with ground and in-flight SRC (Spin Recovery Chute) deploy test as in Figure 14 at February 2004 and finished at December 2005 with approximately 110-sortie flight. Clean, centerline fuel tank, wing tank + SUU20 and each asymmetry configuration was cleared by FSD flight test.

In this section, the T-50/A-50 high AoA flight test results are presented and compared to wind tunnel test results.


8

Figure 14. Ground/In-flight SRC Deployment Test

A. Intentional Departure Test Result The main purposes of intentional departure test are to verify recovery capability and to determine aft CG limit

and to set up recovery procedures. Intentional departure maneuver of T-50/A-50 were performed from low pitch angle and forward CG to high pitch angle and aft CG as a build-up manners.

During T-50/A-50 high AoA test totally over 260 high pitch climbs were attempt and over 230 departures occurred. Intentional departure test began with clean configuration (F0) at forward CG. F0 configuration test was performed as a build up and most high AoA flight test performed with wing-tip missile configurations; Two/one wing-tip missile configurations (F11/F11D1) and two/one wing-tip missile + centerline fuel tank configurations (F10/F10D1).

Figure 15. Two Wing-tip Missile Configuration (F11)

With progress of flight test, the result showed favorable trend than rotary balance wind tunnel test, i.e. inverted deep stall did not occur to the predicted inverted deep stall CG limit. Then test continued to new target CG shifted +1.4%MAC aft considering removal of total 120lbs ballast originally designed for T-50 version to adjust CG. 1. Control Logics inDeparture of T-50/A-50

Control logics in departure of T-50/A-50 are summarized as follow; During an upright pitch departure: Both horizontal tails are commanded to full trailing edge down by the pitch

axis of the FLCS (Flight Control System) to try to reduce the AoA. If AoA exceeds the upper threshold value, the anti-spin logic provides anti-spin commands to the rudder, flaperons, and horizontal tails. Pilot roll and rudder commands are ineffective independent of the MPO (Manual Pitch Override) switch position and pitch stick commands are ineffective without use of the MPO switch.

During an inverted departure: The pitch axis of the FLCS commands the horizontal tails to full trailing edge up to try to return AOA to the normal range. Pitch stick commands are ineffective without use of the MPO switch. If airspeed is low enough or MPO switch is engaged while AoA is below the lower threshold value, the anti-spin logic provides anti-spin commands to the rudder and roll rate feedback commands to flaperons. During an inverted departure, pilot roll and rudder commands should be avoided. Pilot roll and rudder commands are cut out when the MPO switch is engaged.


9

2. Out-of-control Characteristics A departure is a loss of aircraft control that is characterized primarily by uncommanded aircraft motions or

failure of the aircraft’s response to control commands. The FLCS successfully prevented departure, but departures occurred only due to intentional or improper energy management for FSD loading. Departure susceptibility/resistance from normal operation is dealt in section B departure resistance test result.

Inverted deep stall was the biggest concern of high AoA before flight test started. As pre-mentioned, an inverted deep stall was predicted to exist in operational CG range, and speed brakes was estimated effective enough to recover. During FSD flight test, inverted deep stall occurred twice only at most aft CG including some margin of asymmetry missile configuration (F11D1) and aircraft recovered with one cycle of pitch rocking and open SB for each case. In inverted deep stall, aircraft stayed over –50deg AoA, slightly nose-down pitch angle, and over 100KCAS airspeed with generally mild pitch, yaw, and roll rate. If aircraft was not in inverted deep stall, aircraft self-recovered just with release of controls.

Another remarkable characteristics of T-50/A-50 are oscillations in departure. During upright departure, T-50/A-50 showed pitch-yaw-roll coupled oscillation and that oscillation increased with asymmetry configuration. With excessive oscillations, erect hang-up occurred in F10D1 configuration as shown in Figure 16 and recovered with MPO + pitch rocking. The pitch, yaw, and roll rates were over 50/60/100deg/sec during hang-up, and deemed as excessive. The departure transient to upright from inverted was fixed by pitch rate feedback and integrator gain changes and excessive oscillation was reduced by anti-spin logic modification.

-100

-50

0

50

100

AoA

(deg

)

Boom

Selected

WPC+EGI

-40

-20

0

20

40

HT

(deg

)

FTI (Synchro)

Prod.+ Auto Rig

Prod.

-100

-50

0

50

100Rol

l Rat

e, P

B

(deg

/sec

)

FTI (GB002)

Prod. RSA

EGI

-50-30-10103050

0 20 40 60 80 100 120Time (sec)

Yaw

Rat

e, R

B

(deg

/sec

)

FTI (GB003)

Prod. RSA

EGI

-40-20

0204060

Pitc

h St

ick,

FX

(lbs)

Front

SB

AoA Recovery( < 25°)

MPO +Pitch Rocking

SBOpenOvershoot to

Erect Departure

-100

-50

0

50

100

AoA

(deg

)

Boom

Selected

WPC+EGI

-40

-20

0

20

40

HT

(deg

)

FTI (Synchro)

Prod.+ Auto Rig

Prod.

-100

-50

0

50

100Rol

l Rat

e, P

B

(deg

/sec

)

FTI (GB002)

Prod. RSA

EGI

-50-30-10103050

0 20 40 60 80 100 120Time (sec)

Yaw

Rat

e, R

B

(deg

/sec

)

FTI (GB003)

Prod. RSA

EGI

-40-20

0204060

Pitc

h St

ick,

FX

(lbs)

Front

SB

AoA Recovery( < 25°)

MPO +Pitch Rocking

SBOpenOvershoot to

Erect Departure

Figure 16. Erect Hang-up

3. Recovery Characteristics Two/one wing-tip missile configurations (F11/F11D1) self-recovered requiring only release of controls except

two inverted deep stall cases. Self-recovery usually occurred within first two post-departure pitch oscillations. And all recovery took less than 8kft and 30sec to recover including deep stall cases.

Even for inverted deep stall cases, aircraft easily recovered with SB open or MPO + pitch rocking. The open speed bakes generate nose-down pitching moment, which help the aircraft recover. Opening speed brakes was


10

generally enough to break a deep stall, and MPO + pitch rocking was still effective recovery procedure. During MPO + pitch rocking altitude loss was approximately 1.5k to 2k feet per rocking cycle.

Two/one wing-tip missile + centerline fuel tank configurations (F10/F10D1) self-recovered up to most critical operational CG except one erect hang-up case. And it took less than 9kft and 30sec for recovery. Comparing to F11 configuration result, centerline fuel tank made the recovery a little slow. For erect hang-up, near 16kft and 55sec were required to recover with MPO + pitch rocking, but the hang-up was fixed by OFP modifications and did not occur with current OFP.

"F/T": F/T Data , "HQS": Piloted HQS Simulation Data

Pitc

h St

ick

(lbs)

Fx: F/TFx: HQS

AoA

(deg

)

AoA: F/T (Boom)AoA: F/T (Est. EGI)AoA: HQS

HT S

YM

(deg

)

Sym.HT: F/T

0.0 10.0 20.0 30.0 40.0 50.0 60.0 70.0 80.0Time (sec)

KC

AS

(kno

ts)

KCAS: F/T (Boom)KCAS: F/T (Est. EGI)KCAS: HQS

Limiter AoA

Figure 17. Upright Intentional Departure Climb – Departure & Recovery

"F/T": F/T Data , "HQS": Piloted HQS Simulation Data

Pitc

h St

ick

(lbs)

Fx: F/TFx: HQS

AoA

(deg

)

AoA: F/T (Boom)AoA: F/T (Est. EGI)AoA: HQS

HT S

YM

(deg

)

Sym.HT: F/TSym.HT: HQS

0.0 10.0 20.0 30.0 40.0 50.0 60.0 70.0 80.0Time (sec)

KC

AS

(kno

ts)

KCAS: F/T (Boom)KCAS: F/T (Est. EGI)KCAS: HQS

Inverted Departure AoA

Figure 18. Inverted Intentional Departure Climb – Departure & Recovery

Figure 17 and Figure 18 shows typical time history plot of the upright/inverted departure and recovery. The line with symbol shows flight test result and the thick solid line shows HQS (Handling Quality Simulator) result. During pitch climb, airspeed decreased drastically then aircraft was in out of control (departure) by insufficient control power even under AoA-g limiter. But aircraft recovered directly with one AoA overshoot. As plots show, flight test result and HQS result generally matches well.


11

4. Pitching Moment Calculation The pitching moment coefficients were extracted from upright/inverted climb maneuvers for T-50/A-50. The

pitching moment calculated by measured pitch acceleration extracting inertial coupling, engine gyration effect, and engine thrust and drag as following equation.

EngInerAero QQQQ••••

−−= , cSq

IQC

ref

yyAeromAero ××

×=

•

where, ⎩⎨⎧ ××−+−××=

•)()()(1 22

bbxxzzbbzxyy

Iner PRIIPRII

Q

⎭⎬⎫×+×+×+×+

••

)()( bbyxbbxy QPRIRQPI

( )RPMRPM

NINNINIR

Qyy

bEng 2211 ×+××=

•

Figure 19 and Figure 20 shows flight test extracted pitching moment (Cm) with wind tunnel test result of clean configuration (F0) and two wing-tip missile configuration (F11). The plotted pitching moments are with SRC configuration. As the plots show, T-50/A-50 had more nose down tendency for upright and more nose up tendency for inverted in comparison to rotary balance test. And deep stall using rotary balance data deems to be predicted worse than actual. Estimated inverted deep stall CG by flight test data is over 2%MAC aft than rotary balance test prediction and rather similar to low speed wind tunnel data.

0 10 20 30 40 50 60 70 80

AoA, deg

Cm

@ M

RC

35%

MAC

Low Speed W/TRotary Balance W/T Erect 55/CG36 Erect 60/CG36 Erect 65/CG36 Erect 70/CG36 Erect 75/CG36

37%

39%

41%

-80 -70 -60 -50 -40 -30 -20 -10 0

AoA, deg

Cm

@ M

RC

35%

MAC

Low Speed W/TRotary Balance W/T Inverted 70/CG36 Inverted 75/CG36

41%39%37%

Figure 19. Upright/Inverted Flight Test Extracted Cm vs. AoA – F0 Configuration

0 10 20 30 40 50 60 70 80

AoA, deg

Cm

@ M

RC

35%

MAC

Low Speed W/TRotary Balance W/TErect75/CG405Erect75/CG422Erect70/CG422Erect70/CG422Erect70/CG422

37%

39%41%

-80 -70 -60 -50 -40 -30 -20 -10 0

AoA, deg

Cm

@ M

RC

35%

MAC

Low Speed W/T

Rotary Balance W/T

Iverted70/CG422

41%39%37%

Figure 20. Upright/Inverted Flight Test Extracted Cm vs. AoA – F11 Configuration


12

B. Departure Resistance Test Result The purpose of departure resistance test is to verify T-50/A-50’s departure resistance. Most aggressive

maneuvers were performed within flight manual limits. F11/F11D1, F10/F10D1, and F12 were tested. 1. Departrue Prevention Control Logics of T-50/A-50

To prevent departures, several limiting logics were designed and applied to T-50/A-50 OFP; AoA-g limiter, rudder fader, roll rate limiter, high AoA departure prevention logic, and additionally aural warning system.

The T-50/A-50 has automatic AoA-g limiting at both positive and negative AoA in order to automatically prevent pitch axis departures. A rudder fader function helps prevent roll-coupled departures during rudder and rudder-assisted rolls and helps prevent yaw departures during high AoA sideslip maneuvers. The T-50/A-50 roll rate command limit is reduced as AoA & normal load factor increase. 2. Departure Resistance Test Result and Control Law Modification

The T-50/A-50 had no unintended departure during FSD flight test. Some test conditions showed the possibility of departure, but all those conditions was removed by OFP modifications. An excessive overshoot was shown during PA (Power Approach) pushover/pull-up maneuver as shown in Figure 21. That excessive overshoot was fixed by increasing PA pitch rate washout gain.

During UA (Up and Away) negative-g maneuvers in HQS, departure possible test conditions were found. To prevent departure, negative AoA limiter was newly applied to OFP. Figure 22 shows negative-g departure resistance test result of F11/F11D1 configurations with modified OFP. As the plots show, T-50/A-50 showed some AoA overshoot, but no departure.

Figure 23 shows elevated-g departure resistance test result of F11/F11D1. Elevated-g test result was generally stable as shown in the plot and extension of limit AoA could be possible based on that result.

Limiter AoA Overshoot

-20.0

-15.0

-10.0

-5.0

0.0

5.0

10.0

15.0

20.0

100 150 200 250 300

KCAS

(deg

ree)

Min/Max AoS

-10

-8

-5

-3

0

3

5

8

10

100 150 200 250 300

KCAS

(deg

ree)

Min/Max Yaw Rate

-30

-20

-10

0

10

20

100 150 200 250 300

KCAS

(deg

ree/

sec)

Min/Max Roll Rate

-150

-100

-50

0

50

100 150 200 250 300

KCAS

(deg

ree/

sec)

Min/Max AoA

-5

0

5

10

15

20

25

30

35

40

100 150 200 250 300

KCAS

(deg

ree)

δHT

-20

-15

-10

-5

0

5

10

15

20

25

100 150 200 250 300

KCAS

(deg

ree)


-1.00.01.02.03.04.0

100 150 200 250 300 350

KCAS

(deg

ree)F11 30Kft F11D1 30Kft

Removed ByOFP Update


-20.0

-15.0

-10.0

-5.0

0.0

5.0

10.0

15.0

20.0

100 150 200 250 300

KCAS

(deg

ree)

Min/Max AoS

-10

-8

-5

-3

0

3

5

8

10

100 150 200 250 300

KCAS

(deg

ree)

Min/Max Yaw Rate

-30

-20

-10

0

10

20

100 150 200 250 300

KCAS

(deg

ree/

sec)

Min/Max Roll Rate

-150

-100

-50

0

50

100 150 200 250 300

KCAS

(deg

ree/

sec)

Min/Max AoA

-5

0

5

10

15

20

25

30

35

40

100 150 200 250 300

KCAS

(deg

ree)

δHT

-20

-15

-10

-5

0

5

10

15

20

25

100 150 200 250 300

KCAS

(deg

ree)


-1.00.01.02.03.04.0

100 150 200 250 300 350

KCAS

(deg

ree)F11 30Kft F11D1 30Kft

Removed ByOFP Update

Figure 21. Departure Resistance Test Result Summary – PA 1g


13


0.0

5.0

10.0

15.0

100 150 200 250 300 350

KCAS

(deg

ree)


-1.00.01.02.03.04.0

100 150 200 250 300 350

KCAS

(deg

ree)

F11 35Kft F11 20Kft

F11D1 35Kft

Min AoA-20

-15

-10

-5100 150 200 250 300 350

KCAS

(deg

ree)

δHT-20

-15

-10

-5

0100 150 200 250 300 350

KCAS

(deg

ree)

Min/Max AoS

-15

-10

-5

0

5

10

15

100 150 200 250 300 350

KCAS

(deg

ree)

Max Yaw Rate-30

-20

-10

0

10

20

30100 150 200 250 300 350

KCAS

(deg

ree/

sec)

Max Roll Rate

-200

-150

-100

-50

0

50

100

150

200

100 150 200 250 300 350

KCAS

(deg

ree/

sec)


0.0

5.0

10.0

15.0

100 150 200 250 300 350

KCAS

(deg

ree)


-1.00.01.02.03.04.0

100 150 200 250 300 350

KCAS

(deg

ree)

F11 35Kft F11 20Kft

F11D1 35Kft

Min AoA-20

-15

-10

-5100 150 200 250 300 350

KCAS

(deg

ree)

δHT-20

-15

-10

-5

0100 150 200 250 300 350

KCAS

(deg

ree)

Min/Max AoS

-15

-10

-5

0

5

10

15

100 150 200 250 300 350

KCAS

(deg

ree)

Max Yaw Rate-30

-20

-10

0

10

20

30100 150 200 250 300 350

KCAS

(deg

ree/

sec)

Max Roll Rate

-200

-150

-100

-50

0

50

100

150

200

100 150 200 250 300 350

KCAS

(deg

ree/

sec)

Figure 22. Departure Resistance Test Result Summary – UA Negative-g


-1.0

0.0

1.0

2.0

3.0

4.0

100 150 200 250 300 350

KCAS

(deg

ree)

Min/Max AoS

-10

-8

-5

-3

0

3

5

8

10

100 150 200 250 300 350

KCAS

(deg

ree)

Max Yaw Rate

0

10

20

30

40

50

60

100 150 200 250 300 350

KCAS

(deg

ree/

sec)

Max Roll Rate

50

70

90

110

130

150

100 150 200 250 300 350

KCAS

(deg

ree/

sec)


-1.00.01.02.03.04.0

100 150 200 250 300 350

KCAS

(deg

ree)

F11 35Kft F11 20Kft

F11D1 35Kft

Max AoA

20

21

22

23

24

25

26

27

28

100 150 200 250 300 350

KCAS

(deg

ree)

δHT

0

5

10

15

20

100 150 200 250 300 350

KCAS

(deg

ree)

Figure 23. Departure Resistance Test Result Summary – UA Elevated-g

C. Production Configuration Verification Most high AoA tests were performed with SRC configuration as

shown in Figure 24 for safety. In addition to this typical high AoA flight test, final verification with production configuration, i.e. without SRC and noseboom, was planned and performed. It was performed at critical conditions selected from test with SRC configuration. The scope of production configuration verification was same as the SRC configuration test.

The altitude loss and required time to recover of the production configuration are similar to that of the SRC configuration, but the pitch-yaw-roll oscillations of production configuration during erect departure were remarkably reduced. The SRC effects of departure resistance were generally small.

Figure 24. T-50/A-50 SRC Installation


14

V. Conclusion Major design goals of T-50/A-50 high AoA aerodynamics characteristics are the departure resistance, deep stall

prevention, and spin-proof. To achieve the design goals, plenty of low speed and rotary balance wind tunnel tests were planned and

performed. Several analysis and simulation, e.g. deep stall CG limit calculation, deep stall recovery simulation, and departure resistance estimation, were performed using gathered wind tunnel data. During deep stall analysis of T-50/A-50, inverted deep stall condition was found and various resolution options were surveyed at additional wind tunnel test. All static analysis and simulation result using wind tunnel data are presented in this article.

With the foundation of wind tunnel test result, T-50/A-50 flight test was planned and successfully completed with full accomplishment of the design goals. By extracting pitching moment, rotary balance test result was identified too conservative comparing to actual T-50/A-50 high AoA characteristics. With quite good characteristics, test CG was extended more aft from original target, then the ballast, initially planned to be installed on T-50 version, was removed.

Also included is the departure resistance test result. During flight test with initial OFP, a few undesired characteristics were observed, e.g. excessive positive AoA overshoot in PA and negative AoA departure possibility in UA. These potential problems were fixed or improved with proper OFP modifications. With modified new OFP, tests were completed and departure resistance of T-50/A-50 was verified.

In summary, the T-50/A-50, with some OFP modifications, is substantiated to be extremely resistant to departure. The T-50/A-50 will not depart from controlled flight if pilot obey the flight limitations. But T-50/A-50 possibly departs cause by improper energy management or intentional violation of limitations. Even in that departure situation, the T-50/A-50 is proved to have good recovery capability, so that no pilot action is required in operational CG range. Even if inverted deep stall is encountered in case of moving CG beyond current operational range by some reason, to open speed brakes make aircraft recover. In addition, it is recommended to pitch-rock with MPO, because it will be effective and accelerate the recovery.

References 1W. H. Mason, 9. High Angle of Attack Aerodynamics (unidentified document), 2003 2Marilyn E. Ogburn, John V. Foster, Joseph W. Pahle, R. Joe Wilson, and James B. Lackey, Status of the Validation of High-

Angle-of-Attack Nose-Down Pitch Control Margin Design Guidelines (AIAA Paper 93-3623), 1993 3Daniel P. Raymer, Aircraft Design: A Conceptual Approach (AIAA Education Series), 1989 4G. R. Rhodeside, Investigation of aircraft departure susceptibility using a total-G simulator 5Robert Weissman, "Preliminary Criteria for Predicting Departure Characteristics/ Spin Susceptibility of Fighter-Type

Aircraft" (AIAA Paper 86-0492), 1986 6Myung-Sup Lee, Young-Ho Lee, et al., Flight Test Report for High Angle of Attack (85PR0883 TIS FQ-003), 2006 7Lockheed Martin, F-16 S&C Flight Test Report Vol. II (16PR1247), 1979

Documents

2006_6153-T-50-A-50 High Angle of Attack Characteristics